1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to an air cooled turbine airfoil with a TBC or thermal barrier coating.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, be it an aero engine or an industrial gas turbine engine, includes a turbine in which a plurality of stages of stator vanes and rotor blades extract energy from a hot gas flow that passes from the combustor and through the turbine. It is well known in the art of gas turbine engines that the efficiency of the engine can be increased by increasing the hot gas flow entering the turbine. However, the highest temperature obtainable to pass into the turbine is limited to the materials used in the first stage of the stator vane and rotor blades of the turbine.
Providing turbines airfoils (blades and vanes) with cooling air has been used to allow for an increase in the hot gas flow temperature without changing the materials used. Complex internal cooling circuits have been proposed that use convection cooling, impingement cooling and film cooling of the airfoils to prevent over-heating of these airfoils. A turbine airfoil designer wants to provide for maximum cooling of the airfoil while using a minimal amount of cooling air to also increase the efficiency of the engine, since the compressed air used for the internal cooling of the airfoils is typically diverted off from the compressor of the engine. This bleed off air is not used to produce work in the turbine and as such decreases the efficiency of the engine.
Another method of protecting turbine airfoils from extreme heat is to apply a thermal barrier coating (or, TBC) to selective areas of the airfoil that is exposed to the extreme hot temperature. A turbine blade also includes film cooling holes just below the blade tip on both the pressure side wall and the suctions side wall of the blade. The film cooling holes are connected to an internal cooling air supply channel within the blade and are directed to discharge the cooling air upwards and toward the blade tip edge. The TBC is applied on the blade wall from root to tip without covering up the film cooling holes.
The high temperature turbine blade tip section heat load is a function of blade tip leakage flow. A high leakage flow will induce high heat load onto the blade tip section. Thus, blade tip section sealing and cooling have to be addressed as a single problem. Prior art turbine blade tip includes a squealer tip rail which extends around the perimeter of the airfoil and flush with the airfoil wall and forms an inner squealer pocket. The main purpose of incorporating a squealer tip in a blade design is to reduce the blade tip leakage and also to provide for rubbing capability for the blade.
Prior art blade tip cooling is accomplished by drilling holes into the upper extremes of a serpentine flow cooling passage from both of the pressure and suction surfaces near the blade tip edge and the top surface of the squealer cavity. In general, film cooling holes are built into and along the airfoil pressure side and suction side tip sections from the leading edge to the trailing edge in order to provide for edge cooling for the blade squealer tip. Convective cooling holes are also built in along the tip rail at the inner portion of the squealer pocket to provide additional cooling for the squealer tip rail. Since the blade tip region is subject to sever secondary flow leakage field, this translates to a large quality of film cooling holes and cooling flow required in order to adequately cool the blade tip periphery. FIG. 1 shows a prior art turbine blade with a squealer tip cooling arrangement and the secondary hot gas flow migration around the blade tip section. The blade includes a pressure side wall 12 and a suction side wall 13, a squealer pocket 14 formed between a tip rail 15, tip cooling holes 16, and pressure side film cooling holes 17 at the periphery of the tip. A vortex flow 22 from the blade suction side is developed, and a secondary leakage flow 21 flows over the squealer tip. FIGS. 2 and 3 show a profile view of the pressure side and suction side tip peripheral cooling hole configuration for the first stage blade in a turbine. FIG. 2 shows the pressure side tip peripheral film cooling hole pattern with a row of pressure side film cooling holes extending from the leading edge to the trailing edge of the blade. FIG. 3 shows the suction side tip peripheral film cooling hole pattern spaced along the peripheral tip from the leading edge to the trailing edge of the blade. The squealer pocket is formed between the pressure side tip rail and suction side tip rail that extends along the perimeter of the blade tip.
Since the blade squealer tip rail 15 is subject to heating from the three exposed sides—heat load from the airfoil hot gas side surface of the tip rail, heat load from the top portion of the tip rail, and heat load from the back side of the tip rail—cooling of the squealer tip rail by means of a discharge row of film cooling holes along the blade pressure side and suction side peripheral and conduction through the base region of the squealer becomes insufficient. This is primarily due to the combination of squealer pocket geometry and the interaction of the hot gas secondary flow mixing. Thus, the effectiveness induced by the pressure film cooling and tip section convective cooling holes becomes very limited. In addition, a thick TBC is normally used in the industrial gas turbine airfoil for the reduction of the blade metal temperature. However, the TBC is applied around the blade tip rail which may not reduce the blade tip rail metal temperature.
The problem associated with the turbine airfoil tip edge cooling of the prior art can be alleviated by incorporating a new and effective TBC application arrangement of the present invention into the prior art airfoil tip section cooling design.